The present invention relates to an aircraft generator, and in particular to a generator for supplying electrical power to an aircraft when one or more engines thereof has lost propulsive power, or there has been a failure of the main electrical power system.
Aircraft electrical power requirements have increased over the years. This trend is expected to continue as the number of electrical devices, and electrically operated loads increase within aircraft. It is expected that flight control surfaces will increasingly be driven directly by electrical devices, or indirectly, wherein an electrical device is used to provide a supply of hydraulic pressure which may then be used by hydraulic actuators to operate flight control surfaces. With this greater dependence on electrical power, it is becoming more important to ensure that there is an electrical supply available at all times that the aircraft is in service, the main concern is the loss of electrical power in flight in the event of a failure of combustion in the aircraft engines. Such total xe2x80x9cflame-outxe2x80x9d conditions have been known to occur as a result of air turbulence or flying through airborne debris, such as volcanic dust.
Traditionally emergency electrical power in the event of engine failure has been provided by a ram air turbine, RAT, which comprises an electrical generator equipped with a propeller. The RAT is normally stored within the fuselage of an engine and provides no output. However, in emergency conditions, the RAT may be deployed by causing an arm to extend the RAT into the air stream surrounding the aircraft. This flow of air causes the propeller of the RAT to rotate thereby generating electrical power.
It is predicted that it will soon be necessary to provide two RATs on each aircraft in order to ensure sufficient power is available in the event of total engine flame out. This is expected to incur a weight penalty in excess of 250 kg for the RATs plus associated airframe reinforcements.
There are many problems associated with RAT operation. The device itself is a relatively heavy piece of equipment which is carried at all times and which is very rarely deployed, thus it involves a fuel penalty on every flight. Furthermore, the aircraft structure must also be reinforced in the region of the RAT mounting in order to ensure that it can stand the loading experienced in the event that the RAT is deployed. Furthermore, because the RAT is operated only very occasionally and it is not regularly tested for functionality, faults may remain latent for some considerable period of time before being detected. It should be noted that satisfactory operation and deployment of a RAT is not always achieved in practice. It would therefore be advantageous to dispense with the RAT completely.
It is known that, in a multistage high bypass gas turbine engine, the low pressure shaft (LP) or low speed spool which drives the low pressure compressor and the bypass fan will continue to rotate in the event of engine failure because the bypass fan is caused to rotate due to the airflow resulting from the motion of the aircraft as it glides to earth. This is known as the xe2x80x9cwindmillxe2x80x9d effect. The energy of the fan could be extracted by a generator connected to the low pressure shaft which could then supply electrical power to the aircraft during periods of flame-out.
GB 2216603 discloses a gas turbine in which the low speed spool is coupled to an emergency generator via a coupling unit in the event of loss of propulsive drive of the engine. The power take off from the low speed spool may be applied to a gear box which drives an hydraulic pump and an electrical generator. Thus the connection between the low speed spool and the generator is broken when the engine is functioning.
WO93/06007 discloses an arrangement in which the down stream end of the low speed spool is connected to a first gearbox which has an output shaft 26 which extends perpendicularly to the spool and into the interior of the engine pylon. The shaft engages with a second gear box which has various engine accessories coupled to it, such as the main engine fuel pump, a hydraulic pump and one or more electrical generators. This document discloses that the ram air turbine can be dispensed with since power in the event of flame-out conditions can be derived from the windmilling of the bypass fan.
EP 0798454 discloses a multi-spool aero engine in which each of the spools independently and directly drive an electrical generator. The primary source of electrical power is an electric motor/generator positioned within the down stream bearing support structure at the down stream end of the innermost engine shaft. This document goes on to describe that the advantage of this configuration is the elimination of the main engine gearbox and two electrical generators driven by that gearbox. This document also goes on to disclose that the bearings for the shaft may be electromagnetic bearings and that xe2x80x9celectricity generated by the bearing 38 constitutes the primary source of electricity for the aircraft upon which the engine 37 is mountedxe2x80x9d. In order for the magnetic bearing to function as a generator, the magnetic bearing must be directly connected to the shaft, that is there is no intermediate gearbox provided.
EP 0659234 relates to an arrangement in which motor generators are connected to at least two of the engine spools and power transfer can be provided between them. Inductive electrical machines are described as being connected to the low pressure spool via gearing, or alternatively a switched reluctance machine may be provided such that the rotor of the switched reluctance machine is an integral part of the engine shaft, as described in column 10 lines 45 to 50. Thus the switched reluctance machine is directly coupled to the engine shaft.
Although these documents disclose the provision of emergency power by utilising the windmilling of the engine, and two of them disclose providing generators within the structure of the engine itself, none of these documents addresses the issue of obtaining a reasonable amount of power from the generator under the windmilling conditions. The amount of power required should be sufficient to power essential systems until the aircraft descends to an altitude where the auxiliary power unit can be started to initiate an engine restart sequence, and preferably to allow flight control surfaces to be actuated, either directly from an electrical power source or indirectly via an intermediate load transfer system, such as an hydraulic circuit.
According to a first aspect of the present invention, there is provided an electrical generator for use with a gas turbine engine having a low speed spool, comprising a generator driven from a low speed spool, and in which the generator is a switched reluctance generator coupled to the low speed spool via a step up gearbox.
It is thus possible to provide an electrical generator which, by virtue of the step up gearbox, provides a greater electrical output during windmilling conditions of the engine than could be obtained from a similarly sized direct coupled generator.
The maximum available output from a switched reluctance generator is, to a fair approximation, directly related to speed of rotation the mass of the magnetic material forming the rotor and stator of the generator. The step up ratio of the gearbox can be traded against weight of the generator to select a predetermined power output for a given rate of rotation of the low speed spool corresponding to windmilling. Preferably the generator should provide in excess of 10 kW during windmilling, and advantageously should provide around 25 kW or more. However, the step up ratio is limited by the requirement that the generator should still be within safe operating speeds when the low speed spool of the aircraft engine has reached its maximum operating speed. Typically the speed of the low speed spool will vary between 200 and 250 rpm for windmilling and 3000 rpm or so as its maximum operating speed. In an embodiment of the present invention, the gearbox has a step up ratio of 12:1, such that the generator has a rotation rate of around 2400 to 3000 rpm when the engine is windmilling. This enables a generator weighing only approximately 20 kg to produce around 25 kW of power.
The use of a step up gearbox does have the disadvantage of increasing the maximum rotational rate of the generator. In the case of a step up gearbox of 12:1, the maximum rotation rate of the generator is increased to 36,000 rpm or so. It is possible to disconnect the drive to the generator, although this itself has safety implications since the coupling may inadvertently disconnect, or connect.
Preferably the generator is continuously coupled to the low speed spool via the step up gearbox.
In order to be an electrically efficient machine at these high rotational rates, it is desirable that the rotor has a core of thin laminates wherein each laminate is preferably less than 0.5 mm thick. Advantageously the rotor laminate pack is held under compressive load, for example by through bolts. Some of the laminates may be modified so as to define retaining means extending into the inter pole gap between the poles of the laminate core, such that the retaining means engage the bolts holding the laminate core together and serve to restrain the bolts against deformation due to centrifugal force.
Advantageously the stator is also composed of thin laminates.
According to a second aspect of the present invention, there is provided an aircraft mounted gas turbine engine comprising a low speed spool having a plurality of blades at a first end thereof, such that the blades are turned by air flow passing through the engine when a combustion region of the engine is not operating, and a generator connected to the low speed spool, characterised in that the generator is permanently drivingly connected to the low speed spool via a step up gearbox such that the generator can supply power when the low speed spool is being turned by the airflow passing therethrough, and in which the generator is a switched reluctance generator.
Preferably the generator is mounted within the tail portion of the engine.
Preferably the aircraft further includes at least one controller for controlling switching and excitation of a plurality of stator coils within the switched reluctance generator, the controller further being arranged to maintain the output of the generator at substantially the aircraft""s bus voltage. Advantageously, on a multi engine aircraft, each engine is provided with a generator and has an associated generator controller. Each of the generators may then feed directly into an aircraft bus which supplies electrical power to aircraft actuation systems or electro-hydraulic pumps.
According to a third aspect of the present invention, there is provided an aircraft having a plurality of switched reluctance generators, each of which being arranged to deliver power to an aircraft electrical distribution system, in order to supply power to aircraft systems including those for controlling aircraft flight surfaces, wherein the surfaces are directly driven from electrical motors or wherein the flight surfaces are hydraulically operated and the supply of pressurised hydraulic fluid is derived from an electrically driven pump.
According to a further aspect of the present invention, there is provided an aircraft having electrically actuated flight control surfaces and at least one gas turbine engine, and in which the at least one gas turbine engine has a generator permanently drivingly connected to a low speed spool of the engine such that, in the event of flame-out of the at least one engine, the airflow through the engine whilst the aircraft is in flight generates sufficient power to actuate the flight control surfaces.
The flight control surfaces may be linked to an associated control motor via a mechanical linkage, or a motor may be used to drive a pump to supply hydraulic fluid under pressure to hydraulic actuators which in turn actuate the flight control surfaces. It is thus possible to provide an electrically controlled aircraft which can maintain flight control even in the event of complete loss of propulsive power.
The generator may be used only to provide emergency electrical power. In this configuration, the generator will rotate at all times with the low pressure shaft but is only provided with excitation current during engine failure conditions. This simplifies operation of the electrical drive circuits as the drive does not have to be configured to provide excitation at the high operating speeds encountered during normal use. The initial excitation current need only be small and can be provided from a battery or other excitation source eg PM exciter. However, in order to ensure that the emergency generator is capable of operating when required, it may be energised to produce an output during each shut down cycle for the engine. The engine shutdown routine is controlled by a flight management computer which could be arranged to operate the generator as part of the winding down cycle of the engine after each flight. The output of the emergency generator can then be checked against a series of operational parameters and a warning given in the event that the generator fails to operate within the expected range. Alternatively, a confirmation of generator health could be given during each shut down cycle.
Additionally or alternatively, a very low excitation could be provided to the stator windings during all engine operating conditions in order to cause a small monitoring current to be produced. Under such conditions, the generator may have a relatively low output, for example 1 kW or so. This output could then be used to give a continuous, if desired, confirmation of the generator functionality or may be used to flag when a fault condition has occurred.
Given that a multi engine aircraft will carry several emergency generators, redundancy is provided within the generating capability and it is therefore possible that aircraft operation can continue safely after a fault has been detected in one or more of the generators providing that a sufficient number of generators remain operable.
In an alternative mode of operation, the generator may be used continuously at its nominal operating output whilst the aircraft is in use to provide power to the aircraft systems. This does increase the thermal dissipation demands on the generator control electronics and the generator itself, thereby requiring that these components are more robust than would be required if they were to operate in an emergency only mode with or without continuous low power monitoring. However by operating the generator at its nominal full design output, other engine components or aircraft generators may be omitted from the aircraft thereby giving a weight saving.